1. Field of the Invention
The present invention relates generally to gas turbine engines, and more specifically to a turbine stator vane with a cooled leading edge.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine with a number of stages of stator vanes and rotor vanes to guide and extract energy from the hot gas flow passing through the turbine. The first stage vanes and blades are exposed to the highest temperature and therefore are the most critical for cooling. The efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine.
First stage stator vanes used in an industrial gas turbine engine also must be capable of lasting up to 48,000 hours under these extreme operating conditions. Erosion or oxidation damage to a turbine part can cause significant decreases in the engine efficiency or performance. A damaged part, such as the first stage turbine vane, could result in early shutdown of the engine.
In a stator vane, the horseshoe vortex flow phenomenon is created by the combination of hot flow core gas radial velocity and static pressure gradient forces exerted at the intersection of the airfoil leading edge and the endwall of the vane. As the hot flow core gas enters the turbine with a boundary layer thickness and collides with the leading edge of the vane airfoil, the horseshoe vortex separates into a pressure side and suction side downward vortices. Initially, the pressure vortex sweeps downward and flows along the airfoil pressure side forward fillet region first. Then, due to the hot flow channel pressure gradient from the pressure side to the suction side, the pressure side vortex migrates across the hot flow passage and ends up at the suction side of the adjacent airfoil. As the pressure side vortex rolls across the hot flow channel, the size and strength of the passage vortex becomes larger and stronger. Since the passage vortex is much stronger than the suction side vortex, the suction side vortex will flow along the airfoil suction side fillet and act as a counter vortex for the passage vortex.
FIG. 1 shows a schematic view of the vortices formation for a boundary layer entering a turbine airfoil of the prior art. As a result of these vortices flow phenomena, some of the hot core gas flow from the upper airfoil span is transferred toward a close proximity to the endwall and therefore creates a high heat transfer coefficient and high gas temperature region at the airfoil fillet region.
As shown in FIG. 1, the resulting forces drive the stagnated flow that occurs along the airfoil leading edge towards the region of lower pressure at the intersection of the airfoil and endwall. This secondary flow flows around the airfoil leading edge fillet and endwall region. The secondary flow then rolls away from the airfoil leading edge and flows upstream along the endwall against the hot core gas flow as seen in FIG. 2 by reference number 11. As a result, the stagnated flow forces acting on the hot core gas and radial transfer of hot core gas flow from the upper airfoil span toward a close proximity to the endwall creates a high heat transfer coefficient and a high gas temperature at the intersection location.
Injection of film cooling air at discrete locations along the horseshoe vortex region is used to provide the cooling for this region. However, there are many problems for this type of film blowing injection process. High film effectiveness is difficult to establish and maintain in the high turbulent environment and high pressure variation region such as the horseshoe vortex region. Film cooling is very sensitive to pressure gradient. The mainstream pressure variation is very high at the horseshoe vortex location. The spacing between the discrete film cooling holes and areas immediately downstream of the spacing are exposed to less or to no film cooling air at all. Consequently, these areas are more susceptible to thermal degradation and over-temperature.
Cooling of the fillet region by means of conventional backside impingement cooling yields an insufficient result due to the thickness of the airfoil fillet. And, drilling film holes at the airfoil fillet to provide film cooling produces unacceptable stress by the film cooling holes. An alternative way of cooling the fillet region is by injection of film cooling air at discrete locations along the airfoil periphery and endwall into the vortex flow to create a film cooling layer for the fillet region. The film layer migration onto the airfoil fillet region is highly dependant on the secondary flow pressure gradient. For the airfoil pressure side and suction side downstream section, this film injection process provides adequate cooling as seen in FIGS. 4 and 5. However, for the fillet region immediately downstream of the airfoil leading edge, where the mainstream or secondary pressure gradient is in the streamwise direction, the injection of film cooling air from the airfoil or endwall surface will not be able to migrate the cooling flow to the fillet region to create a film sub-boundary layer for cooling that particular section of the fillet.
In addition to the high heat load generated by the horseshoe vortex flow around the airfoil leading edge versus the endwalls junction and the secondary flow accelerated around the airfoil fillet region cooling issue. The first vane normally includes a very long overhang inner diameter endwall and outer diameter endwall on the inlet side of the vane. Use of the long endwalls prevents the use of shallow angled showerhead film cooling holes for the airfoil leading edge region, especially when the showerhead is fabricated with the use of laser drilling for the holes. Drilling holes using a laser requires a larger space between the laser beam and the endwall overhang. As a result of this manufacturing constraint, showerhead hole angles such as 45 degrees are used for the leading edge regions which yields low film coverage, a low film effectiveness and a poor internal convective area. This results in a hotter airfoil leading edge metal temperature or requires more cooling flow for maintain a lower metal temperature that would be provided for by using shallow angled film holes instead.